In aircraft manufacture increasing use is being made, for reasons of weight saving, of composite components which are produced with an epoxy resin composite matrix that is reinforced, in particular, with carbon fibre or glass fibre. Large sized components of composite fibre materials, for example skin panels for aerodynamic effective surface, fuselage barrels or the like, are generally additionally reinforced with stiffening profiles, for example with stringers or annular ribs. The production of stringers and skin panels in the classic CFP/GFP design is expensive because two separate structures are generally required for defining the shape of the CFP material that is initially still soft, followed as a rule by two hardening processes in the autoclave. Furthermore, lost cores of hard foam or the like have to be used in many cases for predefining the shape. These cores remain in the component after the hardening process and result in an increase in weight without performing an additional reinforcing function. Moreover, serious tolerance problems arise because two cured components have to be glued together to provide an exact fit in order to maintain a predetermined theoretical contour within the predetermined tolerances. In this procedure equalising the tolerances between the components is only possible by means of the adhesive applied in the contact area and initially still flexible.
Although an alternative method of production, in which the stringers or the skin panel have not yet fully cured when joined together, generally requires one further passage through the autoclave, it incurs an additional positioning and upgrade cost in terms of the shape defining structure in order to keep the parts to be joined together in their theoretical position within the close tolerances generally specified, and avoid any relative displacements.